The present invention relates generally to aircraft landing gear, more particularly to semi-levered landing gear, and most particularly to a control system for controlling a landing gear such that semi-levered functionality is achieved at selected operating conditions.
In most large commercial aircraft, the maximum rotation angle of the aircraft during takeoff and landing is limited by a minimum permissible clearance between a rear under portion of the fuselage and the ground. It is known that the takeoff and landing performance of a given aircraft can be enhanced by providing a longer main landing gear about which the aircraft rotates to achieve a nose-up attitude, thereby increasing the maximum rotation angle of the aircraft. However, one of the objectives of aircraft design is to configure the landing gear so that the aircraft fuselage is essentially horizontal during ground operations and has an appropriate sill height for ground servicing. The maximum sill height that is acceptable is dictated by the height of ground equipment that must interface with the aircraft, and thus is generally fixed. In many cases, the maximum allowable sill height is less than what would be desirable from an aircraft performance standpoint, and therefore, merely lengthening the landing gear is not a viable approach to achieving increased maximum rotation angle. Further, landing gear length must be minimized to keep weight to a minimum and to facilitate the stowing of the gear during flight, and hence a wholesale lengthening of the landing gear is undesirable.
In view of the above considerations, efforts have been made to develop variable-length landing gear capable of assuming a length that is suitable for stowing within the aircraft, and for ground operations while the aircraft is on the ground and stationary, and further capable of assuming a greater length during takeoff and landing operations. One such type of variable-length landing gear, to which the present invention relates, is the semi-levered landing gear (SLG). In a typical SLG, a wheel truck is formed by a bogie beam supporting forward and aft wheels at forward and aft ends thereof, and a main strut of conventional design is pivotally connected to the bogie beam at a main pivot between the forward and aft wheels. An additional mechanical linkage is connected at an upper end to the main strut and at a lower end to the bogie beam at an auxiliary pivot spaced from the main pivot for controlling positioning of the bogie beam. The additional mechanical linkage enables the bogie beam, under certain conditions, to pivot about the auxiliary pivot rather than the main pivot. In this manner, when the aircraft approaches the end of a takeoff roll and begins to rotate for liftoff, the bogie beam can be placed in a tilted orientation with the forward wheels off the ground with the aid of the additional mechanical linkage, which prevents the bogie beam from rotating to a horizontal orientation. With the wheel truck in this tilted position, the effective length of the landing gear is increased relative to its length when all wheels are on the ground. The aircraft can then rotate to a higher pitch attitude, with the same tail clearance, thus achieving improved takeoff performance.
Existing semi-levered landing gears can be unsatisfactory for various reasons. In some types of SLG configurations, such as that disclosed in U.S. Pat. No. 4,892,270 to Derrien et al., the additional mechanical linkage comprises a passive torque link assembly whose only function is to lock up when the main strut and the bogie beam assume particular positions, namely, when the bogie beam is tilted and the main strut is relatively uncompressed as it is on initial touchdown and at liftoff. These types of SLG devices require an additional actuator or spring device for placing the bogie beam in the tilted position for landing. Where the means for tilting the bogie beam is a passive spring device as in the Derrien ""270 patent, stowing of the landing gear in the aircraft can be complicated by the lack of ability to reposition the bogie beam in a more-appropriate position for stowage.
One method that has been used to reposition the bogie for stowage with this type of SLG employs a shrink-link main strut that is operable to shorten as the landing gear is retracted into the wheel well, thereby changing the geometry of the SLG link and bogie so that the gear can be stowed. A disadvantage of this approach is that the shrink-link main strut is of considerably greater complexity and weight than a conventional main strut, thereby adding cost and weight to the aircraft.
Accordingly, some SLG configurations employ an active device connected between the main strut and the bogie beam for placing the bogie beam in a tilted position. For example, published UK Patent Application No. GB 2,101,542A by Putnam et al. discloses an aircraft undercarriage unit having a variable length oleopneumatic strut connected between the main strut and an aft end of the bogie beam. The variable length strut is hydraulically actuated to extend so as to tilt the bogie beam during takeoff. After takeoff, the variable length strut is contracted to position the bogie beam substantially horizontal to facilitate stowage of the gear. A major problem with Putnam""s landing gear design is that it is incapable of maintaining equal loading on all main gear wheels during braking at all aircraft weight and aerodynamic lift conditions, because the variable-length strut is always active to exert a force on the bogie tending to tilt the bogie, which occurs when the overall load on the landing gear drops to a sufficiently low level. The result is that Putnam""s landing gear would require larger brakes, and larger wheel wells to contain them, in order to assure adequate braking capacity during landing rollout or refused takeoff, thus incurring a significant penalty to the aircraft design in terms of weight and wheel well volume.
Another type of main landing gear is disclosed in UK Patent 1,510,554 by Faithfull. The Faithfull patent states as its object and advantage the capability of effectively lengthening the landing gear at touchdown to provide improved shock absorbing characteristics during landing at relatively high descent rates. The landing gear purportedly achieves this object by the use of an additional oil-filled cylinder, functioning only as a passive damper, pivotally attached to the front of the bogie beam and the upper stationary part of the main shock strut. In preparation for landing, the bogie is placed into a tilted position via a positioning device that is separate from the oil-filled cylinder. In this tilted position, the oil-filled cylinder is in a compressed condition. Upon touchdown and landing rollout, the bogie begins to rotate toward a horizontal position, thus causing the oil-filled cylinder to be extended until it reaches its maximum length. The maximum length of the oil-filled cylinder is such that the bogie cannot rotate to a fully horizontal position during the initial portion of the landing rollout, and hence the effective length of the landing gear is greater during this initial portion of the rollout.
Faithfull does not claim that his device is capable of providing improved takeoff performance through effective gear lengthening. Moreover, Faithfull""s device would prevent the most advantageous positioning of the bogie for stowage of the gear in the aircraft. In order to stow the landing gear in most aircraft, the bogie advantageously should be placed in an approximately horizontal position (on some large commercial aircraft, the bogie must rotate past horizontal into a pitch-down attitude of as much as 15 degrees) with the main strut fully extended, this orientation enabling the wheel well size to be kept to a minimum. However, Faithfull""s oil-filled cylinder has a maximum extension selected such that the bogie is tilted into a pitch-up attitude when the main strut is slightly compressed on landing. Thus, the oil-filled cylinder simply cannot extend sufficiently to position the bogie horizontal with the main strut fully extended. If the oil-filled cylinder disclosed in Faithfull were modified to provide sufficient stroke to accommodate the bogie stow position, it would be incapable of providing the semi-levered function on landing. Furthermore, if the stroke length were selected to provide effective semi-levered function on takeoff, then the bogie would assume a pitch-up attitude for stowage, which would require a very large wheel well. Thus, Faithfull""s device is incapable of simultaneously providing semi-levered function and enabling an optimum positioning of the bogie for stowage.
A main landing gear configuration disclosed in U.S. Pat. No. 4,749,152 is said to provide an effectively longer landing gear at takeoff, but requires a very complex main strut having multiple main strut cylinders, some with offset loading. This main strut would result in a very heavy landing gear relative to a conventional main strut. Additionally, the landing gear in the ""152 patent requires a shrink-link mechanism to reposition the bogie for stowage. Furthermore, the multiple-cylinder design results in sliding surfaces that cannot be inspected without major disassembly, thus increasing maintenance costs. Finally, another disadvantage of the gear design disclosed in the ""152 patent is that all of the purported functions of the gear, including semi-levered action at takeoff, absorption of energy at touchdown, equal wheel loading during ground roll, and bogie repositioning, are provided by the main strut. This may hamper the optimization of each of these functions because of space and geometry limitations of the design.
To address the above-noted needs, the assignee of the present application developed a landing gear with an auxiliary strut as described in U.S. Pat. No. 6,182,925, the disclosure of which is hereby incorporated herein by reference. The ""925 patent describes a semi-levered landing gear that includes a single auxiliary strut in conjunction with a main strut, which can be of conventional design, and a multiple-wheeled bogie. The auxiliary strut, in preferred embodiments, enables the landing gear to provide all of the desirable functions required of a main gear during aircraft operation, including:
(1) the ability to tilt the bogie to provide an effectively longer main landing gear during takeoff rotation and liftoff;
(2) the ability to reposition the bogie beam to an appropriate angle for stowing the landing gear;
(3) the ability to position the bogie beam to an appropriate pitch-up angle in preparation for landing after landing gear deployment, and thereby facilitate an early air-ground sensing upon initial ground contact of the aft bogie wheels;
(4) the ability to effectively decouple the auxiliary strut during static and ground-roll operations so as to facilitate equal loading of all main gear wheels and, accordingly, optimum braking ability; and
(5) the ability to deactivate the functioning of the auxiliary strut that provides the semi-levered action when desired, such as during landing, so that the auxiliary strut acts as a damping device for partially absorbing touchdown loads such that the load transmitted to the aircraft is reduced.
To these ends, the semi-levered landing gear of the ""925 patent comprises a wheel truck including a bogie beam and at least one forward wheel and at least one aft wheel rotatably supported by the bogie beam at forward and aft portions thereof, respectively, a main strut having an upper portion and a lower portion telescopingly connected to each other such that the main strut is extendable and compressible, the lower portion having a lower end pivotally connected to the bogie beam at a main pivot located between the forward and aft wheels, and an auxiliary strut having an upper end pivotally connected to the upper portion of the main strut and a lower end pivotally connected to the bogie beam at an auxiliary pivot longitudinally spaced from the main pivot. The auxiliary strut comprises a cylinder barrel having a closed end and an open end, a piston assembly slidably received through the open end of the cylinder barrel, and a lock-up device operable to permit extension of the piston assembly during a portion of a stroke thereof until the auxiliary strut reaches a predetermined lock-up length between a maximum length and a minimum length thereof. The lock-up device substantially prevents further extension of the piston assembly once the auxiliary strut reaches the predetermined lock-up length. The main strut and auxiliary strut are constructed and arranged relative to each other and the bogie beam such that, during takeoff as the main strut extends, the auxiliary strut becomes locked at the predetermined lock-up length before the main strut fully extends such that further extension of the main strut causes the bogie beam to pivot about the auxiliary pivot so as to tilt the bogie beam, whereby the landing gear is effectively lengthened. The main strut can be of conventional design; no shrink-link or other complex and heavy main strut is needed.
The present invention represents a further development of the technology embodied in the ""925 patent. More particularly, the present invention relates to a control system and method for controlling a tiltable wheel truck of a main landing gear by employing an auxiliary strut that can be commanded to lock up at a predetermined length so as to cause the landing gear to function as a semi-levered gear. The control system and method in particular operate to cause the landing gear to function as a semi-levered gear during a takeoff roll; the auxiliary strut remains unlocked during other operating conditions so that, for example, the wheel truck is free to pivot during rollout following a landing, and during taxi such that the truck can pitch freely in response to runway surface roughness. The auxiliary strut is unlocked at touchdown also, so that the auxiliary strut can provide a damping function. Preferably, the auxiliary strut and control system can also provide a truck positioning function for placing the truck in a desired orientation. For example, the truck can be tilted into a pitched-up attitude for touchdown so that the aft wheels of the gear make first contact with the runway; furthermore, the truck can be pivoted to an orientation suitable for stowing the gear in the gear bay of the aircraft.
In accordance with one embodiment of the invention, a control system for controlling a tiltable wheel truck of a main landing gear includes an auxiliary strut, a ground mode sensor operably connected with the aircraft for detecting and providing signals indicative of when the aircraft is on the ground; a takeoff mode sensor operably connected with the aircraft for detecting and providing signals indicative of when the aircraft is operating in a throttled-up mode; and an auxiliary strut control unit operably connected with the ground mode sensor, takeoff mode sensor, and auxiliary strut. The auxiliary strut control unit is operable to issue a lock-up command signal to the auxiliary strut upon detecting signals from the sensors indicating that the aircraft is on the ground and that the aircraft is operating in a throttled-up mode, whereby the auxiliary strut is caused to lock up during a takeoff roll but is unlocked during other operating modes of the aircraft.
Preferably, the ground mode sensor comprises a weight-on-wheels sensor that detects when the main landing gear is bearing weight, thus indicating that the landing gear is in contact with the ground. The takeoff mode sensor preferably comprises an engine speed sensor operable to detect when any of the aircraft""s engines is operating above a predetermined speed. This allows the control unit to distinguish between a takeoff condition at which the engines will be operating at a relatively high speed (e.g., greater than 60 percent fan speed) and a taxi condition or landing rollout at which the engines typically operate at a relatively low speed (e.g., less than 60 percent fan speed).
In order for the control unit to provide an immediate unlock signal to the auxiliary strut in the event of a refused takeoff (RTO), the takeoff mode sensor preferably also comprises a thrust lever sensor operable to detect when any of the thrust levers for the aircraft""s engines is advanced beyond a predetermined limit, which indicates a throttled-up condition. Thus, if an RTO occurs during a takeoff roll and the thrust levers are chopped back to idle (i.e., below the predetermined limit), the auxiliary strut is immediately unlocked so that the load on the landing gear is evenly distributed to all wheels for maximum braking efficiency.
The takeoff mode sensor preferably also comprises a ground speed sensor operable to detect when the aircraft is traveling above a predetermined ground speed. In this manner, the control unit is able to distinguish between a takeoff roll and, for instance, a ground test of the engines in which the engines may be operating at a high speed. Thus, the auxiliary strut can remain unlocked unless the aircraft is actually rolling down the runway at an appreciable speed.
Preferably, the control unit is operable to unlock the auxiliary strut upon expiration of a predetermined time period following liftoff of the aircraft from the ground. Liftoff is indicated by a change of state of the signal from the weight-on-wheels sensor. Additionally or alternatively, if desired, the control unit can unlock the auxiliary strut immediately upon detecting that a command has been issued to retract the landing gear.
In accordance with another preferred embodiment of the invention, detection of whether the aircraft is on the ground or in the air is accomplished through use of radio altimeters mounted in the aircraft. Commercial aircraft typically have at least one and more typically three radio altimeters mounted in the aircraft at a location on the underside of the fuselage, usually just behind the nose gear. Such a radio altimeter emits radio signals downwardly, and the signals bounce off the ground and are reflected back up to the aircraft. The altimeter receives the reflected signal and computes the height of the altimeter from the ground based on the time delay between the sent and received signals. The altimeter is usually calibrated so that it reads zero when the main landing gear of the aircraft just touch down on landing, at which point the nose gear of the aircraft where the altimeter is located is still some distance off the ground. Thus, when the nose gear touches down and the aircraft is in a landing roll, the altimeter will read a negative height, about xe2x88x928 to xe2x88x9210 feet, for example. In the preferred embodiment of the invention, the signal(s) from one or more radio altimeters is (are) used to determine whether the aircraft is on the ground or in the air.
For example, where there are three radio altimeters for redundancy, the control logic determines that the aircraft is on the ground if any of the following conditions is satisfied for a predetermined length of time (e.g., 5 seconds): (1) data from all three altimeters are available and any two altimeters indicate a height less than or equal to a predetermined lower limit (e.g., xe2x88x925 feet); or (2) data from one or two altimeters are unavailable but at least one altimeter indicates a height less than or equal to the predetermined lower limit; or (3) data from all altimeters are unavailable. The on-ground condition remains true until the in-air logic is satisfied.
The control logic determines that the aircraft is in the air if any of the following conditions is satisfied for a predetermined length of time (e.g., 5 seconds): (1) data from all three altimeters are available and any two altimeters indicate a height greater than a predetermined upper limit (e.g., 15 feet); or (2) data from one or two altimeters are unavailable, and all available altimeters indicate a height greater than the predetermined lower limit and at least one altimeter indicates a height greater than the predetermined upper limit. The in-air condition remains true until the on-ground logic is satisfied.
In accordance with this embodiment, the auxiliary strut is commanded to lock if each engine is running, and flaps are in the takeoff position for more than a predetermined length of time (e.g., 2 seconds), and the aircraft is on the ground, and any engine power lever is advanced past a predetermined limit (e.g., 40 degrees). The command to lock remains until the unlock logic is satisfied.
The auxiliary strut is commanded to unlock if any of the following conditions persists for more than a predetermined length of time (e.g., 1 second): (1) the aircraft is in the air; or (2) each engine power lever is pulled back below the predetermined limit (e.g., 40 degrees) and the aircraft ground speed is greater than a predetermined limit (e.g., 60 knots); or (3) any engine is not running and the aircraft ground speed is less than the predetermined limit (e.g., 60 knots); or (4) all engines are not running; or (5) flaps are not in the takeoff position. The unlock command remains until the lock logic is satisfied, even if some of the above conditions are no longer satisfied.
This control logic provides robustness so that various failure modes can be tolerated while maintaining appropriate operation of the auxiliary strut.